An example high-performance system may include an example high-performance component. The high-performance component may include a substrate defining a channel. The channel defines a leading ramp and a trailing ramp. The example high-performance component includes an abradable track between the leading and the trailing ramps. The abradable track includes a porous abradable composition. The example high-performance system may include a rotating component configured to contact and abrade the abradable track. An example technique for forming the abradable track includes thermal spraying a precursor composition at the channel to form the abradable track.

In one exemplary embodiment, an airfoil includes pressure and suction side walls that extend in a chord-wise direction between leading and trailing edges. The pressure and suction side walls also extend in a radial direction to provide an exterior airfoil surface. A cooling passage is arranged between the pressure and suction walls. The cooling passage has a first width in the chord-wise direction near the suction side wall. A second width is in the chord-wise direction near the pressure side wall. A third width is between the pressure and suction side walls. The third width is smaller than the first and second widths.

An apparatus and method relating to a cooling hole of a component of a turbine engine. The component can include a wall separating the hot gas fluid flow from the cooling fluid flow and having a heated surface along which the hot gas fluid flow flows and a cooled surface facing the cooling fluid flow and at least one cooling hole comprising at least one inlet at the cooled surface, at least one outlet at the heated surface, with the outlet having a modified outlet shape.

A pressure-release device has a pressure-release flap arranged in a delimiting wall of a chamber of an aircraft engine which opens in the presence of a predefined pressure in the chamber. The pressure-release flap includes a flap region and a hinge. The flap region is defined by a perforation formed in the delimiting wall which delimits the flap region with respect to a surrounding region of the delimiting wall. The perforation is formed such that, in the presence of the predefined pressure in the chamber, the flap region pivots open relative to the surrounding region, with the perforation rupturing, and the hinge is fixedly connected to the surrounding region and the flap region, such that the hinge holds the flap region on the surrounding region in the event of a rupture of the perforation and a pivoting-open of the flap region.

A blade assembly for use in a gas turbine engine. The blade assembly includes a blade, a platform distinct from the blade and configured to extend around the blade, and a pin that couples the platform with the blade.

A method of controlling an extent of a thermal barrier coating (TBC) sheet spall and a hot gas path (HGP) component are disclosed. The method provides an HGP component having a body with an exterior surface. Controlling the extent of the TBC sheet spall includes forming a TBC over a selected portion of the exterior surface of the body. The TBC includes a plurality of segments in a cellular pattern. Each segment is defined by one or more slots in the TBC, and each segment has a predefined area such that the extent of the TBC sheet spall is limited by the predefined area of each of the plurality of segments that constitute the TBC sheet spall.

A seal device for a turbine is disposed around a rotor so as to separate a high-pressure space and a low-pressure space and includes: a plurality of thin plates arranged along an outer peripheral surface of the rotor. Each of the thin plates has a thin-plate tip surface facing the outer peripheral surface of the rotor; a first side plate disposed so as to face the high-pressure space and covering outer peripheral regions of first side surfaces; and a second side plate disposed so as to face the low-pressure space and covering outer peripheral regions of second side surfaces. The first side surface of each of the thin plates is covered with the first side plate in a region extending further to an inner side, in a radial direction of the rotor, than a region of the second side surface covered with the second side plate.

Disclosed is a tangential on-board injector (TOBI) system that includes an annulus and a plurality of cooling airflow passages disposed about the annulus. Each cooling airflow passage of the plurality of cooling airflow passages includes an inlet opening having a polygonal inlet cross-section, the inlet opening having an inlet cross-sectional area. Each cooling airflow passage of the plurality of cooling airflow passages further includes an outlet opening having an outlet cross-section and an outlet cross-sectional area. The inlet cross-sectional area is greater in magnitude than the outlet cross-sectional area. Also disclosed are additive manufacturing methods for manufacturing the tangential on-board injector system and gas turbine engines that incorporate the tangential on-board injector system.

An apparatus and method relating to a cooling hole of a component of a turbine engine. The component can include a wall separating the hot gas fluid flow from the cooling fluid flow and having a heated surface along which the hot gas fluid flow flows and a cooled surface facing the cooling fluid flow and at least one cooling hole comprising at least one inlet at the cooled surface, at least one outlet at the heated surface, with the outlet having a modified outlet shape.

A gas turbine engine component includes a passage and a wall adjacent the passage. The wall includes a first side bordering the passage and a second side opposite the first side. The second side includes an array of cells. The wall also includes an array of channels. Each of the channels is located proximate a corresponding one of the cells. A coating is disposed over the cells. When the coating degrades the channels open to permit impingement air flow through the channel onto sidewalls of the cells.